Rocket / Thruster
Resistojet Thruster Pulse Combustion Hall Thruster Deflagration Thruster Mono-Fuel Thruster
Testing Log
Gaseous-Fuel Rocket
Nitrous oxide mono-fuel thruster Overall view of the nitrous oxide monofuel thruster test setup View through the nitrous oxide monofuel thruster nozzle of heated catalyst LabVIEW data acquisition control panel for the nitrous oxide monofuel thruster hot test number three

Nitrous Oxide Mono-Fuel Thruster

Jet Aerospace designed, analyzed, fabricated, and tested this thruster in September 2005. The decomposition of a mono-fuel, like hydrogen peroxide or nitrous oxide, might be a good choice for small impulse thrusters. Although the specific impulse will be considerably less than that of popular hypergolic combinations like monomethyl hydrazine and nitrogen tetroxide, the processing simplicity of propellants like nitrous oxide might outweigh performance limitations.

The application for a thruster of this size would be for small satellite attitude adjustment. As the intent of the thruster is to fire in small intermittent bursts, there is no integrated cooling system. The thruster was tested with a fixed mass, high pressure gas charge. The gas was heated in the injection system and passed through a metal oxide catalyst bed. Heat is generated in the catalyst from the decomposition of the nitrous oxide into nitrogen and oxygen. The performance of the thruster and transient behavior of the catalyst was tested by measuring temperature change across the catalyst and pressure drop across the nozzle.

The testing program consisted of propellant cold-flow and thruster hot-fire testing. The cold-flow tests were used as a benchmark for a no-reaction condition. Hot-fire tests were run with several injector temperatures and catalyst preheat purge conditions. The thruster produced an average specific impulse of 157 seconds at 0.14 pounds of thrust with pulse duration of 0.6 seconds. Although the specific impulse of the overall thruster is relatively high, some portion of the impulse was provided by resistojet type thrust from injector heating. However, the nozzle inlet temperature profile suggests some level of N2O decomposition and heat generation in the catalyst bed. Maximum gas injection temperature was measured at 678 degrees Fahrenheit at the beginning of gas pulse. The calculated nozzle throat velocity was somewhat lower than expected with a Mach number of 0.61 (with respect to local conditions). The lower than expected nozzle throat velocity shows that the nozzle throat sizing was somewhat optimistic based on pretest flow calculations. The exhaust gas did not luminesce in any of the tests.

Specifications

Weight

     

2.5 lb

Chamber Dimensions

     

7.45" L / 1.65" D

Propellant

     

Pulsed Gaseous Nitrous Oxide (N2O)

Reaction Chamber

     

Insulated Pressure Vessel

Chamber Cooling

     

None

Nozzle

     

Convergent / Divergent

Reaction Initiation

     

Resistance Heated Propellant Injection
Purged Catalyst Heating

Instrumentation

     

Temperature, Pressure

Project Information

Years(s) Built

     

2005

Number of Test Runs

     

12 Cold Tests
24 Hot Tests

Intent of Design

     

To test performance of nitrous oxide mono-fuel thruster

Testing Results

     

Maximum Specific Impulse of 157 seconds
Maximum thrust of 0.14 lb
Maximum nozzle Mach number of 0.61
Showed favorable nozzle pressure increase in hot tests
Nozzle pressure pulse duration - 0.6 seconds

jetman@jetaerospace.org